A gas turbine engine such as a turbofan engine used for powering an aircraft in flight typically includes in serial flow communication about an axial centerline axis a fan, compressor, combustor, high pressure turbine (HPT), and a low pressure turbine (LPT). Compressed air is discharged from the compressor into the combustor wherein it is mixed with fuel and ignited for generating combustion gases which flow downstream through the HPT and the LPT. The HPT is conventionally joined to the compressor and extracts energy from the combustion gases for powering the compressor. And, the LPT is conventionally joined to the fan and extracts energy from the combustion gases for powering the fan.
In order to protect engine components from the hot combustion gases, a portion of the compressed air is bled from the compressor and channeled to the components for their cooling. However, any air so bled from the compressor decreases the overall efficiency of the engine since it is not being used to generate the combustion gases from which energy is extracted by the turbines. The cooling air is discharged from the components after providing cooling thereof and rejoins the combustion gases flowing through the engine which creates additional efficiency losses due to the mixing therewith. Accordingly, it is desirable to minimize the amount of cooling air bled from the compressor for improving overall efficiency of the engine while obtaining acceptable cooling of the hot components.
Since the HPT firstly receives the combustion gases from the combustor, it is subject to the hottest combustion gas temperatures and requires suitable cooling. The HPT typically includes one or more rotor stages with a stationary nozzle disposed upstream of each rotor stage. The stage-one nozzle is disposed at the outlet of the combustor and first receives the combustion gases therefrom which are suitably channeled by the nozzle into the stage-one rotor blades. The nozzle includes radially outer and inner bands between which are fixedly joined a plurality of circumferentially spaced apart nozzle vanes. The vanes are typically hollow and provided with compressor bleed air for cooling the vanes, with another portion of the bleed air being suitably channeled to both the outer and inner bands for their cooling as well. Since the outer and inner bands extend for a substantial axial distance, they require suitable cooling along their entire axial length.
For example, the nozzle inner band typically includes a plurality of circumferentially spaced apart, conventional film cooling holes extending radially therethrough which are disposed upstream of the throats between the adjacent nozzle vanes. These film cooling holes are provided to form a cooling air film which extends downstream therefrom to the aft edge of the inner band for providing acceptable cooling thereof. The film cooling holes are provided upstream of the nozzle throats to minimize mixing losses thereof with the combustion gases flowing axially between the vanes. However, this places the film cooling holes substantially upstream from the aft edge of the inner band which requires a suitable quantity of cooling air for ensuring effective cooling from the film cooling holes axially downstream to the inner band aft edge. This configuration also typically results in overcooling of the inner band intermediate region wherein the film cooling holes discharge the cooling air, with a decreasing amount of cooling axially along the inner band to the aft edge thereof. Accordingly, axial thermal gradients are created in the inner band between the film cooling holes and the aft edge which results in undesirable thermal stress therein.